Nuclear thermal propulsion rocket engine

ABSTRACT

A fission based nuclear thermal propulsion rocket engine. An embodiment provides a source of fissionable material such as plutonium in a carrier gas such as deuterium. A neutron source is provided, such as from a neutron beam generator. By way of engine design geometry, various embodiments may provide for intersection of neutrons with the fissionable material injected by way of the carrier gas, while in a reactor provided in the form of a reaction chamber. Impact of neutrons on fissionable material results in a nuclear fission in sub-critical mass reaction conditions in the reactor, resulting in release of heat energy to the materials within the reactor. The reactor is sized and shaped to receive the reactants and an expandable fluid such as hydrogen, and to confine heated and pressurized gases for discharge out through a throat, into a rocket engine expansion nozzle for propulsive discharge.

COPYRIGHT RIGHTS IN THE DRAWING

A portion of the disclosure of this patent document contains materialthat is subject to copyright protection. The patent owner has noobjection to the facsimile reproduction by anyone of the patent documentor the patent disclosure, as it appears in the Patent and TrademarkOffice patent file or records, but otherwise reserves all copyrightrights whatsoever.

RELATED PATENT APPLICATIONS

None.

TECHNICAL FIELD

This disclosure relates to rocket engines, and more specifically, torocket engines which utilize nuclear fission as the source for thermalenergy in the creation of motive force to create specific impulsesufficient for lifting objects to earth orbit, or for insertion intointerplanetary flight.

BACKGROUND

A continuing interest exists for improvements in rocket engines, andmore particularly for designs that would provide a significant increasein propulsive power, as often characterized by the benchmark of specificimpulse, especially as might be compared to conventional chemicallyfueled rocket engines. Such new rocket engines might be useful in avariety of applications. Launch operational costs might be substantiallyreduced on a per pound of payload basis, by adoption of a new nuclearthermal propulsion rocket engine design that provides significantimprovements in the specific impulse, as compared to existing prior artrocket engine designs. Further, from the point of view of overallmission costs, since the mass of most components of rocket vehicles areproportional to the mass of the propellant, it would be desirable todevelop a new rocket engine design that reduces the mass of consumablecomponents necessary for initiating lift off and acceleration to orbitalvelocity. Such an improvement would have a major impact on the entirefield of rocket science from a launch weight to payload ratio basis.Such an efficiency improvement would also facilitate the inclusion ofwings and recovery systems that would enable an economic fully reusablelaunch system with airliner type operations, for example, as describedin U.S. Pat. No. 4,802,639, issued Feb. 7, 1989 to Richard Hardy et al.,entitled HORIZONTAL TAKEOFF TRANSATMOSPHERIC LAUNCH SYSTEM, thedisclosure of which is incorporated herein in its entirety by thisreference. And, for missions beyond earth orbit, it would beadvantageous, from the point of view of mission duration, to provide anew rocket engine design that reduces not only the payload to launchweight, but also the transit time to the mission objective, by providinghigh specific impulse, so as to minimize fuel required to achieve highvehicle velocities necessary to accomplish a selected interplanetarymission in a specific time frame. And, it would be desirable to providesuch an improved rocket engine that includes components which have beenreused and identified as comparatively reliable and cost effective, andthus, minimizes design risk and thus minimizes the extent of testingthat may be necessary, as compared to many alternate designs which aresubject to stress and strain from temperature and pressure in rocketengine services. Thus, it can be appreciated that it would beadvantageous to provide a new, high efficiency rocket engine designwhich provides a high specific impulse, thus minimizing the launchweight to payload ratio.

In general, the efficiency of a rocket engine may be evaluated by theeffective use of the consumable propellant, i.e. the amount of impulseproduced per mass of propellant, which is itself proportional to thevelocity of the gases leaving the rocket engine nozzle. In nuclearthermal rocket engine systems, the specific impulse increases as thesquare root of the temperature, and inversely as the square root of themolecular mass of the gases leaving the rocket engine nozzle.Consequently, in the design of a nuclear thermal rocket engine,efficiency is maximized by using the highest temperature available,given materials design constraints, and by utilizing a fluid that has avery low molecular mass for generation of thrust.

A variety of fission based rocket engines have been contemplated, andsome have been tested. An overview of the current status of suchefforts, and suggestions as to suitable configurations for variousmissions, was published on Oct. 16, 2014, at the Angelo State UniversityPhysics Colloquium in San Angelo, N. Mex., by Michael G. Houts, Ph.D, ofthe NASA Marshall Space Flight Center, Huntsville, Ala., in hispresentation entitled Space Nuclear Power and Propulsion; available atwebsite: ntrs.nasa.gov/search.jsp?R=20140016814. As he notes, theRover/NERVA program (1955-1973) tested a fission rocket engine design.Further, the most powerful nuclear rocket engine that has been tested,to date, was the Phoebus 2a, which utilized a reactor that was operatedat a power level of more than 4.0 million kilowatts, during 12 minutesof a 32 minute test firing. However, it is clear that the variousnuclear fission rocket engine designs currently available have variousdrawbacks, such as excessive gamma radiation production of retained corecomponents, which requires extensive and heavy shielding, if used onmanned missions.

One of the more interesting disclosures of a fission based rocket enginewas provided in U.S. Pat. No. 6,876,714 B2, issued on Apr. 5, 2005 toCarlo Rubbia, which is titled DEVICE FOR HEATING GAS FROM A THIN LAYEROF NUCLEAR FUEL, AND SPACE ENGINE INCORPORATING SUCH DEVICE, thedisclosure of which is incorporated herein in its entirety by thisreference. That patent discloses the heating of hydrogen gas by fissionfragments emitted from a thin film of fissile material, such asAmericium metal or a compound thereof, which is deposited on an innerwall of a cooled chamber. However, that device generally describes theuse of fissile material in critical mass conditions, and although itmentions the contemplation of sub-critical mass fission arrangements,details of such a condition are scant, if indeed present at all in thedescription thereof.

Thus, there remains a need to provide a design for a high specificimpulse nuclear thermal propulsion rocket engine that simultaneouslyresolves two or more of the various practical problems, including (a)minimizing the weight of consumables (such as chemical fuelconstituents) on a per payload pound basis; (b) avoiding excessiveradiation shielding requirements when the design is used in mannedmissions, by avoiding use of retained radioactive hardware thatgenerates large gamma ray emissions; (c) providing for power control,especially as related to power generation amounts at any given time, byproviding for throttling of the fission reaction; and (d) providing ahigh specific impulse, as compared to prior art rocket engines for spacevehicles.

SUMMARY

A novel fission based nuclear thermal propulsion rocket engine has beendeveloped, which, in various embodiments, significantly enhances thespecific impulse provided by the propulsion system. The rocket enginedesign provides source of fissionable material such as plutonium in acarrier gas such as deuterium. A neutron source is provided from aneutron beam generator. By way of engine design geometry, variousembodiments may provide for intersection of a neutron beam from theneutron generator with the fissionable material injected by way of acarrier gas into a reactor provided in the form of a reaction chamber.Impact of the neutron beam on the fissionable material results in anuclear fission in sub-critical mass reaction conditions in the reactor,resulting in release of heat energy to the materials within the reactor.The reactor is sized and shaped to receive the reactants and to receivean expandable fluid such as hydrogen, and to confine heated andpressurized gases for discharge out through a throat, into a rocketengine expansion nozzle for propulsive discharge therefrom.

An advantage of the novel fission based nuclear thermal propulsionrocket engine design disclosed herein is that use of such a design in asecond stage of a launch system would result in all radioactive fissionproducts being exhausted into the vacuum of space.

Moreover, recent developments in neutron beam generators has madepossible the development of a nuclear thermal rocket engine in which theprocess of production of neutrons can be partially separated from theprocess of absorption of neutrons by fissionable material, so that thefission process can be initiated and maintained while utilizing lessthan a critical mass of fissionable material. In this manner, a designhas been developed in which radioactive fission products ejected out ofthe rocket nozzle into space with other exhaust gases, while amounts offissionable material consumed are replenished with new fissionablematerial only as necessary to support continued fission, to obtain thenecessary heat release for operation.

BRIEF DESCRIPTION OF THE DRAWING

The present invention(s) will be described by way of exemplaryembodiments, using for illustration the accompanying drawing in whichlike reference numerals denote like elements, and in which:

FIG. 1 is a partial cross sectional view for an embodiment of a rocketengine, showing a reactor in the form of a reaction chamber with arestrictive throat forming an outlet which leads to an expansion nozzle,and a neutron beam generator provides a beam of neutrons into thereactor to intersect with actinides injected into the reactor with afirst fluid, and also showing injection of a second fluid which isprovided to provide thrust for expansion due to heating in the reactor,as well as diagrammatically depicting turbopumps for providing both afirst fluid and a second fluid to the reactor.

FIG. 2 is a partial perspective view for an embodiment of a rocketengine, showing the components just mentioned with respect to FIG. 1above showing use of a neutron beam generator that provides a beam ofneutrons into the reaction chamber to intersect with actinides injectedinto the reaction chamber with a first fluid, and also showing injectionof a second fluid which is provided to provide thrust for expansion dueto heating in the reaction chamber, as well as conceptually depictingturbopumps for providing both a first fluid and a second fluid to thereaction chamber, and also showing use of a gas generator for developinghigh pressure combustion gases for driving a fuel turbopump and a thrustfluid turbopump.

FIG. 3 is a top view of an embodiment for a reaction chamber, showing alocation for a neutron beam generator, and also showing coolantpassageways which run along the outer surface of the sides and top ofthe reaction chamber to a header which collects the heated second fluidand from which the second fluid is injected into the reaction chamber.

FIG. 4 is a bottom view of a an embodiment for a rocket engine, takenlooking up at line 4-4 of FIG. 1, showing the second fluid distributorat the outlet of the expansion nozzle that is used to distribute thesecond fluid to coolant passageways along the walls of the expansionnozzle and the reaction chamber, and also showing the outlet of thereaction chamber.

FIG. 5 is a perspective view of an embodiment for a rocket engine,showing a neutron beam generator mounted to a reaction chamber, and anexpansion nozzle for receiving heated gases from the reaction chamber,as well as showing coolant passageways on outer surfaces of the reactionchamber and on the outer surfaces of the expansion nozzle.

FIG. 6 is similar to FIG. 1 above, showing a partial cross sectionalview for an embodiment of an rocket engine, depicting a reaction chamberwith a restrictive throat forming an outlet which leads to an expansionnozzle, and a neutron beam generator provides a beam of neutrons intothe reaction chamber to intersect with actinides injected into thereaction chamber with a first fluid, and also showing injection of asecond fluid which is provided to provide thrust for expansion due toheating in the reaction chamber, as well as diagrammatically depictinguse of turbopumps on a common shaft or gear box for providing both afirst fluid and a second fluid to the reaction chamber, and also having,driven by the same turbine system, an electrical generator forgenerating electrical power for supply to the neutron beam generator.

The foregoing figures, being merely exemplary, contain various elementsthat may be present or omitted from a final configuration for anembodiment of a nuclear thermal rocket engine using sub-critical massfission of fuels, or that may be implemented in various embodimentsdescribed herein for a rocket engine. Other variations in nuclearthermal rocket engine designs may use slightly different mechanicalstructures, mechanical arrangements, solid flow configurations, liquidflow configurations, or vapor flow configurations, and yet employ theprinciples described herein and as generally depicted in the drawingfigures provided. An attempt has been made to draw the figures in a waythat illustrates at least those elements that are significant for anunderstanding of exemplary nuclear thermal rocket engine designs undersub-critical mass fission conditions. Such details may be quite usefulfor providing propulsion for a high specific impulse space vehicle, andthus, reduce cost of payloads lifted to earth orbit, lunar, orinterplanetary missions.

It should be understood that various features may be utilized in accordwith the teachings hereof, as may be useful in different embodiments asuseful for various sizes and shapes, and thrust requirements, dependingupon the mission requirements, within the scope and coverage of theteachings herein as defined by the claims. Further, like features invarious nuclear thermal rocket engine designs may be described usinglike reference numerals, or other like references, without furthermention thereof.

DETAILED DESCRIPTION

Attention is directed to FIGS. 1 and 2 of the drawing. FIG. 1 shows in apartial cross sectional view an embodiment for a nuclear thermal rocketengine 10, showing a reactor 12 having a tubular portion 13 and arestrictive throat 14 forming an outlet 16 which leads to an expansionnozzle 20. A neutron beam generator 22 is provided to direct a beam ofneutrons 24 in the reactor 12. A first fluid storage compartment 26 forstorage of a first fluid 28 such as deuterium (may be depicted as ₁D² oras ²H) is provided. A second fluid storage compartment 30 is providedfor storage of a second fluid 32 such as hydrogen H₂. A third fluidstorage compartment 34 is provided for storage of third fluid 36 such asoxygen (O₂). In an embodiment, the third fluid 36 may be used forreaction with a second fluid 32 such as hydrogen (H₂) in a gas generator38 (also marked as GG in FIGS. 1 and 2), to generate a high pressurefluid 40 (e.g., combustion gases) for driving a turbine 42 in aturbopump 44, as further discussed below. In such case, after pressurereduction through turbine 42, remaining low pressure water vapor may bedischarged overboard as indicated by reference arrow 46.

A selected actinide fuel F which provides a fissile material may besupplied from storage container 50 for mixing with the first fluid 28.In an embodiment, a selected fuel F may be provided in a particulateform. In an embodiment, the selected fuel F may be provided in a veryfine particulate, or more specifically, in a finely powdered form. In anembodiment, the powdered fuel may comprise a selected actinide compound.In an embodiment, the powdered fuel may comprise a substantially puremetallic actinide. In an embodiment, the fuel F may be supplied in aform including of one or more plutonium (Pu) isotopes. In an embodiment,the fuel F may be supplied in as a fissile material in the form ofplutonium 239 (²³⁹Pu). In an embodiment, the fuel F may be supplied as afissile material in the form of uranium 235 (²³⁵U). In variousembodiments, the selected fissile material providing fuel F, beforeinjection into the reactor 12, may be provided in particulate form.

In an embodiment, the first fluid 28 from the first fluid storagecompartment 26 may be mixed with a selected amount of fuel F, beforeinjection into reactor 12. In an embodiment, the first fluid 28 and aselected amount of fuel F may be mixed to create a rich fuel mixture 52,before passage of the rich fuel mixture 52 (i.e. a mixture of fuel F andfirst fluid 28) through control valve 53 and then into a fuel turbopump54, which pumps the fuel rich mixture 52 into reactor 12 via fuel supplyline 56, fuel header 58, and a first set of fuel injectors 60 whichconfine and direct passage of fuel rich mixture 52 into reactor 12. Inan embodiment, control valve 53 may provide on-off capability. Invarious embodiments, control valve 53 may additionally providethrottling capability to regulate the quantity of flow of the rich fuelmixture 52. In an embodiment, at time of injection, the fuel richmixture 52 may be in gaseous form, while carrying a particulate actinidefuel F therein. However, as shown in FIGS. 1 and 2, in variousembodiments, at time of injection, the fuel rich mixture 52 may be inliquid form, while carrying a particulate actinide fuel F therein. In anembodiment, a first set of fuel injectors 60 may be oriented at aselected inwardly directed angle alpha (α) that directs a rich fuelmixture 52 stream toward a reaction zone 62 wherein energetic neutrons24 from neutron beam generator 22 collide with atoms of fissile materialin fuel F as found in the rich fuel mixture 52, to cause fission ofatoms of fuel F, with resultant heat release. In any event, a neutronbeam generator 22, which is further discussed below, is configured todirect neutrons 24 to collide with at least some of the fuel F fissilematerial in the reaction zone 62, wherein the neutrons 24 and thefissile material interact to thereby effect fission of at least some ofthe atoms of the fissile material in fuel F and release heat.

In various embodiments, a rocket engine 10 may operate with fission ofthe fissile material of fuel F under sub-critical mass conditions. Undervarious embodiments, the fissile material may include plutonium 239. Inan embodiment the amount of plutonium 239 (²³⁹Pu) provided may bebetween about thirty parts per million (30 ppm) and about one hundredand twenty parts per million (120 ppm), by weight, in the first fluid28. In an embodiment the amount of plutonium 239 (²³⁹Pu) provided may bebetween about sixty parts per million (60 ppm) and ninety parts permillion (90 ppm), by weight, in the first fluid 28. In an embodiment ofrocket engine 10, plutonium 239 (²³⁹Pu) may be provided at about sixtyparts per million (60 ppm), by weight, in the first fluid 28. In variousembodiments, the first fluid 28 may be provided as deuterium (may beshown as either ₁D² or ²H).

In various embodiments, the first fluid 28 may include one or moreisotopes of hydrogen. In an embodiment, the first fluid 28 may includedeuterium. In an embodiment, the first fluid 28 may primarily bedeuterium (²H). In an embodiment the first fluid 28 may includeessentially only deuterium (₁D²). In an embodiment, the first fluid 28may include at least some tritium (₁T³). In an embodiment, the firstfluid 28 may include both deuterium and tritium. In an embodiment, thepresence of tritium may induce secondary fusion reactions in the centerof the fluid flow while being directed out through the nozzle, therebyincreasing specific impulse without significantly increasing engine walltemperature.

To provide thrust, by way of heating and expansion in the reactor 12 andresultant expulsion out thru expansion nozzle 20, a low molecular weightfluid such as hydrogen (H₂) is provided as the second fluid 32. A secondfluid 32 may be stored in a second fluid storage compartment 30, and ondemand is delivered by line 70 to the thrust fluid turbopump 44. Thethrust fluid turbopump 44 receives the second fluid 32 from the secondfluid storage compartment 30 and provides (generally indirectly) thesecond fluid 32 under pressure to the reaction chamber 12. In anembodiment, the second fluid 32 may be send under pressure from thrustfluid turbopump 44 via second fluid supply line 72 to a distributionring 74 located at or near the exit plane 77 of expansion nozzle 20. Thesecond fluid 32 may be supplied via distribution ring 74 to nozzlecoolant passageways 76 located on the exterior 78 of expansion nozzle20. In this manner, an extremely cold fluid, e.g. liquid hydrogen, maybe utilized as a coolant for the expansion nozzle by passage of thesecond fluid 32 through the nozzle coolant passageways 76. Likewise, asalso seen in FIG. 3, the reactor 12 includes reactor coolant passageways86 on the reactor external surface 88. In this manner, an extremely coldfluid, e.g. liquid hydrogen, is utilized as a coolant for the reactor 12by passage of the second fluid 32 through the reactor coolantpassageways 86. Thus, the rocket engine 10 may utilize the second fluid32 as a coolant by way of the passage of second fluid 32 through thenozzle coolant passageways 76 and through the reactor coolantpassageways 86, before injection of the second fluid 32 into the reactor12.

Once second fluid 32 reaches the upper end 90 of reactor 12, acollection header 92 may be utilized to accumulate the second fluid 32from the reactor coolant passageways 86. In an embodiment, fromcollection header 92, the second fluid 32 may be directed to a secondset of injectors 94 which are configured for confining the passage ofthe second fluid 32 during injection into the reactor 12. By way ofinjectors 94, the second fluid 32 may be directed toward or injectedinto a mixing zone 96, which mixing zone 96 is located downstream of thereaction zone 62. In mixing zone 96, the second fluid 32 is heated andexpanded, in order to provide thrust by ejection through throat 14 andoutlet 16 of reactor 12. Also, the first fluid 28 is heated andexpanded, in order to provide thrust by ejection through throat 14 andoutlet 16 of reactor 12.

As mentioned above, in order to provide power for the thrust fluidturbopump 44, a gas generating chamber 38 may be provided to generatecombustion products in the form of a hot gas 40 that drives a turbine42, which in turn drives a pump impeller 100. Consequently, when oxygen36 is supplied for combustion with hydrogen as second fluid 32, watervapor is formed, and the resultant low pressure water vapor stream 46 isdischarged overboard. Likewise, hydrogen as second fluid 32 and oxygen36 may be supplied to a second gas generating unit 102 to generate hotgas 104 that drives turbine 106 which in turn drives fuel pump impeller108 in fuel turbopump 54.

In another embodiment for rocket engine 10′ as seen in FIG. 6, adifferent design for a thrust fuel turbopump 144 may be provided. Insuch design, the thrust fuel turbopump 144 may provide pumping of secondfluid 32 by pump impeller 145, while also additionally providing anelectrical generator 146. In an embodiment, the electrical generator 146may be configured to generate electrical power, and supply the same viaelectrical power lines 148 and 150 to neutron beam generator 22. In anembodiment a thrust fluid turbopump 144 may further include a fuelturbopump 160, for receiving first fluid 28 from the first fluid storagecompartment 26 and providing the first fluid 28 under pressure toreactor 12. In an embodiment, the thrust fluid turbopump rotor 145, thefuel turbopump rotor 161, and the electrical generator 146 may all bedriven by a gas turbine 162 on a common shaft 164 or via gearbox from acommon shaft 164.

In various embodiments for a rocket engine 10 or 10′ or the like, usingnuclear thermal heating of a low molecular weight gas such as hydrogenas described herein, a rocket engine may be provided that has a specificimpulse in the range of from about 800 to about 2500 seconds. In variousembodiments using nuclear thermal heating of a low molecular weight gassuch as hydrogen as described herein, a rocket engine may be providedthat has a specific impulse in the range of from about 1000 to about1215 seconds.

To summarize, in order to facilitate supply of hydrogen to the reactor12 for heating, a thrust fluid turbopump 44 or 144 or the like may beprovided as generally described herein above. In an embodiment, liquidhydrogen, i.e. a cryogenic liquid, may be provided to the rocket engine10 or 10′, by way of a thrust fluid turbopump that is driven by aturbine which is rotatably energized by high temperature gases. In anembodiment, the high temperature gases may be provided by way ofcombustion products, such as by way of combustion of hydrogen and oxygenin a gas generating chamber GG to generate a high temperature combustiongas, which after passage through the turbine 42 or 162, as the case maybe, may be exhausted overboard in the form of a water vapor stream 46 or46′. The tradeoff of loss of efficiency due to loss of propellant(hydrogen) expended in the gas generating chamber GG, in view of theusual weight savings and simplicity of design (and lack of radioactivecontamination), as compared to additional weight and complexity in viewof any additional specific impulse contribution in designs that mightavoid such combustion losses, may be evaluated for a specific spacevehicle design and attendant mission profile, as will be understood bythose of skill in the art. Various configurations for drive of asuitable thrust fluid turbopump for feeding hydrogen to the reactionchamber may be provided by those of skill in the art using conventionalliquid turbopump system design principles, and thus, it is unnecessaryto provide such details. In general, the thrust fluid turbopump mustavoid cavitation while pumping liquid hydrogen at relatively low inletpressure, and deliver the hydrogen to the reaction chamber (and in anembodiment, via distribution ring and cooling passageways) at very highpressure, and preferably, with capability to provide a relatively widethrottling range. In various embodiments, the selected thrust fluidturbopump 44 or 144 design may be optimized for minimizing weight whileproviding necessary performance while at the same time minimizing thethrust fluid turbopump package size, in order to minimize necessaryspace in a selected space vehicle design. Selection of suitable bearingsand seals are of course necessary, and various design alternatives areknown to those of skill in the art. More generally, those of skill inthe art will understand that turbopumps for supply of cryogenic liquidsto rocket engines require designs that provide maximum performance atminimum weight.

Similarly, to facilitate supply of the plutonium carrying deuterium gasto the reactor 12 for fission of at least some of the plutonium, a fuelturbopump 54 may be provided. In an embodiment, liquid deuterium i.e. acryogenic liquid, may be provided to the rocket engine 10 or 10′, by wayof a fuel turbopump 54 or 160, that is driven by a turbine (106 or 162)which is rotatably energized by high temperature gases. In anembodiment, the high temperature gases may be provided by way ofcombustion products, such as by way of combustion of hydrogen and oxygento generate a high temperature combustion gas. Various configurationsfor drive of a suitable fuel turbopump for feeding deuterium (andplutonium carried therewith) to the reaction chamber may be provided bythose of skill in the art using conventional liquid turbopump systemdesign principles, and thus, it is unnecessary to provide such details.In general, the fuel turbopump (54 or 160) must avoid cavitation whilepumping liquid deuterium at relatively low inlet pressure, and deliverthe deuterium to the reaction chamber at very high pressure, andpreferably, with capability to provide a relatively wide throttlingrange. In various embodiments, the selected fuel turbopump design may beoptimized for minimizing weight while providing necessary performancewhile at the same time minimizing fuel turbopump package size, in orderto minimize necessary space in a selected space vehicle design.

Further, in order to generate electricity for a selected neutron beamgenerator 22, an electrical generator 146 may be combined with aturbopump 144, so that a hot gas driven turbine 162 in the turbopump 144also provides shaft power for an electrical generator 146. In anembodiment, the high temperature gases may be provided by way ofcombustion products, such as by way of combustion of hydrogen and oxygenin a gas generating chamber GG to generate a high temperature combustiongas, which after passage through the gas turbine 162, may be exhaustedoverboard via a water vapor exhaust tube 46. Alternately, a stand-aloneelectrical turbine generator may be provided, with its own hydrogen gasor combustion gas driven turbine, in the manner as generally describedabove.

In an embodiment, a deuterium-deuterium (“DD”) type neutron generator 22may be utilized. As an example, high yield neutron generators arecurrently available for various applications with variable neutronoutput between 1×10¹¹ and 5×10¹¹ neutrons per second (n/s). It is anadvantage of a DD type neutron generator design that because no tritiumis utilized, radiation shielding and accompanying safety concerns andregulatory burdens are significantly reduced. Thus, such designs may bemore suitable for manned space vehicles.

However, in an embodiment, a deuterium-tritium (“DT”) type neutrongenerator may be utilized. As an example, extremely high yield neutrongenerators based on DT design principles are currently available withvariable neutron output between 1×10¹³ and 5×10¹³ neutrons per second(n/s). Such designs may require appropriate shielding and regulatoryapprovals for manned spaceflight applications, but may be especiallysuitable for high payload unmanned spaceflight vehicle applications.

Neutron generators of either deuterium-deuterium design or ofdeuterium-tritium design have been developed by Phoenix Nuclear Labs,2555 Industrial Drive, Monona, Wis. 53713, with a web page atphoenixnuclearlabs.com. Other vendors currently provide differentdesigns. For example, Gradel Group, 6, Z.A.E. Triangle Vert, L-5691ELLANGE, Luxembourg (see website atgradel.lu/en/activities/neutrons-generators/products/14-1-mev-neutrons-dt/)currently provides a 14 MeV neutron generator of deuterium-tritiumdesign, with basic functionality as follows:₁ D ²+₁ T ³→₂ He ⁴(3.5 MeV)+₀ n ¹(14.1 MeV)

It is currently anticipated that any selected neutron beam generatordesign may require adaptive configurations to various structures andcomponents to make them suitable for the rigors of a rocket launch andsubsequent spaceflight environment. However, the fundamental principlesdescribed herein for creation of a fission based rocket engine may beachieved by provision of a suitably adapted neutron beam generatordevice.

In the foregoing description, for purposes of explanation, numerousdetails have been set forth in order to provide a thorough understandingof the disclosed exemplary embodiments for the design of a nuclearthermal rocket engine operable in sub-critical mass fissile conditions.However, certain of the described details may not be required in orderto provide useful embodiments, or to practice selected or otherdisclosed embodiments. Further, for descriptive purposes, variousrelative terms may be used. Terms that are relative only to a point ofreference are not meant to be interpreted as absolute limitations, butare instead included in the foregoing description to facilitateunderstanding of the various aspects of the disclosed embodiments. And,various actions or activities in any method described herein may havebeen described as multiple discrete activities, in turn, in a mannerthat is most helpful in understanding the present invention. However,the order of description should not be construed as to imply that suchactivities are necessarily order dependent. In particular, certainoperations may not necessarily need to be performed precisely in theorder of presentation. And, in different embodiments of the invention,one or more activities may be performed simultaneously, or eliminated inpart or in whole while other activities may be added. Also, the readerwill note that the phrase “in an embodiment” or “in one embodiment” hasbeen used repeatedly. This phrase generally does not refer to the sameembodiment; however, it may. Finally, the terms “comprising”, “having”and “including” should be considered synonymous, unless the contextdictates otherwise.

It will be understood by persons skilled in the art that variousembodiments for novel nuclear thermal rocket engine designs utilizingsub-critical mass fission of a selected actinide fissile material havebeen described herein only to an extent appropriate for such skilledpersons to make and use such nuclear thermal rocket engine. Additionaldetails may be worked out by those of skill in the art for a selectedset of mission requirements and design criteria, such as whether themission is manned or unmanned, (e.g., whether any necessary radiationminimization or radiation shielding may be required). Although onlycertain specific embodiments of the present invention have been shownand described, there is no intent to limit this invention by theseembodiments. Rather, the invention is to be defined by the appendedclaims and their equivalents when taken in combination with thedescription.

Importantly, the aspects and embodiments described and claimed hereinmay be modified from those shown without materially departing from thenovel teachings and advantages provided, and may be embodied in otherspecific forms without departing from the spirit or essentialcharacteristics thereof. Therefore, the embodiments presented herein areto be considered in all respects as illustrative and not restrictive orlimiting. As such, this disclosure is intended to cover the structuresdescribed herein and not only structural equivalents thereof, but alsoequivalent structures.

Numerous modifications and variations are possible in light of the aboveteachings. Therefore, the protection afforded to this invention shouldbe limited only by the claims set forth herein, and the legalequivalents thereof.

We claim:
 1. A rocket engine, comprising: a neutron beam generator; afirst fluid storage compartment for storage of a first fluid; a secondfluid storage compartment for storage of a second fluid; a reactor, saidreactor comprising (a) a chamber for containing said first fluid andsaid second fluid during heating, (b) a first set of injectors for (i)confining passage into said reactor of said first fluid received fromsaid first fluid storage compartment, said first fluid containing atleast some fissile material, and (ii) for injecting said first fluidcontaining at least some fissile material to a reaction zone in saidreactor, (c) a second set of injectors for directing passage into saidreactor of said second fluid received from said second fluid storagecompartment; and (d) an outlet; an expansion nozzle, said expansionnozzle connected to said outlet of said reactor; wherein said neutronbeam generator is configured to direct neutrons to collide with at leastsome of said fissile material in said reaction zone, wherein saidneutrons and said fissile material interact to thereby effect fission ofat least some of said fissile material and release heat; and whereinsaid first fluid and said second fluid are contained by and heated insaid reaction chamber to produce a heated gas which is released throughsaid outlet and then expelled through said expansion nozzle.
 2. Therocket engine as set forth in claim 1, wherein said first fluidcomprises one or more isotopes of hydrogen.
 3. The rocket engine as setforth in claim 1, wherein said first fluid comprises deuterium.
 4. Therocket engine as set forth in claim 1, wherein said fissile materialcomprises an actinide.
 5. The rocket engine as set forth in claim 1,wherein said fissile material comprises one or more Pu isotopes.
 6. Therocket engine as set forth in claim 5, wherein said fissile materialcomprises plutonium 239 (²³⁹Pu).
 7. The rocket engine as set forth inclaim 4, wherein said actinide comprises uranium 235 (²³⁵U).
 8. Therocket engine as set forth in any one of claim 4, 5, 6, or 7, whereinsaid fissile material, before injection into said reactor, is providedin particulate form.
 9. The rocket engine as set forth in claim 1,wherein fission of said fissile material occurs under sub-critical massconditions, and wherein said fissile material comprises ²³⁹Pu, andwherein said ²³⁹Pu is provided at about between thirty parts per millionand one hundred and twenty parts per million, by weight, in said firstfluid.
 10. The rocket engine as set forth in claim 1, wherein fission ofsaid fissile material occurs under sub-critical mass conditions, andwherein said fissile material comprises ²³⁹Pu, and wherein said ²³⁹Pu isprovided at about between sixty parts per million, and ninety parts permillion, by weight, in said first fluid.
 11. The rocket engine as setforth in claim 8, wherein said first fluid comprises a gas at time ofinjection into said reactor.
 12. The rocket engine as set forth in claim1, wherein said expansion nozzle comprises nozzle coolant passageways.13. The rocket engine as set forth in claim 12, wherein said secondfluid is utilized as a coolant by passage through said nozzle coolantpassageways, before injection of said second fluid into said reactor.14. The rocket engine as set forth in claim 13, wherein said secondfluid, at time of entry into said nozzle coolant passageway, comprises aliquid.
 15. The rocket engine as set forth in claim 1, wherein saidreactor comprises reactor coolant passageways.
 16. The rocket engine asset forth in claim 15, wherein said second fluid is utilized as acoolant by passage through said reactor coolant passageways, beforeinjection of said second fluid into said reactor.
 17. The rocket engineas set forth in claim 16, wherein said second fluid, at time of entryinto said reactor coolant passageways, comprises a liquid.
 18. Therocket engine as set forth in claim 1, wherein said second fluid isinjected into said reactor at a mixing zone, said mixing zone locateddownstream of said reaction zone.
 19. The rocket engine as set forth inclaim 1, wherein said reaction chamber comprises a tubular shapedportion.
 20. The rocket engine as set forth in claim 1, wherein saidrocket engine has a specific impulse in the range of from about 800 toabout 2500 seconds.
 21. The rocket engine as set forth in claim 1,wherein said rocket engine has a specific impulse in the range of fromabout 1000 to about 1215 seconds.
 22. A rocket engine, comprising: aneutron beam generator; a first fluid storage compartment for storage ofa first fluid, and wherein said first fluid comprises deuterium ₁D²; asecond fluid storage compartment for storage of a second fluid, whereinsaid second fluid, before heating, comprises hydrogen (H₂); a reactor,said reactor comprising (a) a chamber for containing said first fluidand said second fluid during heating, (b) a first set of injectors for(i) directing passage into said reactor of said first fluid receivedfrom said first fluid storage compartment, said first fluid containingat least some fissile material, wherein said fissile material comprisesat least some ²³⁹Pu, and (ii) for injecting said first fluid containingat least some ²³⁹Pu into a reaction zone in said reactor, (c) a secondset of injectors for directing passage into said reactor of said secondfluid received from said second fluid storage compartment; (d) saidreactor further comprising a mixing zone, and wherein said second fluidis injected into said reactor at said mixing zone, said mixing zonelocated downstream of said reaction zone, and (e) an outlet; anexpansion nozzle, said expansion nozzle connected to said outlet of saidreactor; wherein said neutron beam generator is configured to directneutrons to collide with at least some of said fissile material in saidreaction zone, wherein said neutrons and said fissile material interactat a common point in said reactor, to thereby effect fission of at leastsome of said fissile material at sub-critical mass conditions, andrelease heat therefrom; and wherein said first fluid and said secondfluid are contained by and heated in said reaction chamber to produce aheated gas which is released through said outlet and expelled throughsaid expansion nozzle.
 23. The rocket engine as set forth in claim 22,wherein said ²³⁹Pu, before injection into said reactor, is provided inparticulate form.
 24. The rocket engine as set forth in claim 22,wherein said ²³⁹Pu is provided at about sixty parts per million, byweight, in said deuterium (₁D²).
 25. The rocket engine as set forth inclaim 22, wherein said first fluid comprises a gas at time of injectioninto said reactor.
 26. The rocket engine as set forth in claim 22,wherein said expansion nozzle comprises nozzle coolant passageways. 27.The rocket engine as set forth in claim 26, wherein said second fluid isutilized as a coolant by passage through said nozzle coolantpassageways, before injection into said reactor.
 28. The rocket engineas set forth in claim 26, wherein said second fluid, at time of entryinto said nozzle coolant passageway, comprises a liquid.
 29. The rocketengine as set forth in claim 22, wherein said reactor comprises reactorcoolant passageways.
 30. The rocket engine as set forth in claim 29,wherein said second fluid is utilized as a coolant by passage throughsaid reactor coolant passageways, before injection into said reactor.31. The rocket engine as set forth in claim 30, wherein said secondfluid, at time of entry into said reactor coolant passageways, comprisesa liquid.
 32. The rocket engine as set forth in claim 1, or in claim 22,further comprising a fuel turbopump, said fuel turbopump receiving saidfirst fluid from said first fluid storage compartment, and providingsaid first fluid under pressure to said reaction chamber.
 33. The rocketengine as set forth in claim 1, or in claim 22, wherein said first fluidfurther comprises tritium (₁T³).
 34. The rocket engine as set forth inclaim 32, further comprising an oxygen storage compartment and a gasgenerating chamber, and wherein said fuel turbopump is driven bycombustion products formed by combustion of hydrogen and oxygen in saidgas generating chamber.
 35. The rocket engine as set forth in claim 1,or in claim 22, further comprising a thrust fluid turbopump, said thrustfluid turbopump receiving said second fluid from said second fluidstorage compartment, and providing said second fluid under pressure tosaid reaction chamber.
 36. The rocket engine as set forth in claim 35,further comprising an oxygen storage compartment and a gas generatingchamber, and wherein said thrust fluid turbopump is driven by combustionproducts formed by combustion of hydrogen and oxygen in said gasgenerating chamber.
 37. The rocket engine as set forth in claim 35,wherein said thrust fluid turbopump further comprises an electricalgenerator, said electrical generator configured to generate electricalpower, and to supply electrical power to said neutron beam generator.38. The rocket engine as set forth in claim 35, wherein said thrustfluid turbopump further comprises a fuel turbopump, said fuel turbopumpreceiving said first fluid from said first fluid storage compartment,and providing said first fluid under pressure to said reaction chamber.39. The rocket engine as set forth in claim 35, wherein said thrustfluid turbopump further comprises an electrical generator, saidelectrical generator configured to generate electrical power, and tosupply electrical power to said neutron beam generator and wherein saidthrust fluid turbopump further comprises a fuel turbopump, said fuelturbopump receiving said first fluid from said first fluid storagecompartment, and providing said first fluid under pressure to saidreaction chamber, and wherein said thrust fluid turbopump, said fuelturbopump, and said electrical generator are all driven by a gas turbineon a common shaft or via gearbox from a common shaft.
 40. The rocketengine as set forth in claim 1, or in claim 22, wherein fission of saidfissile material occurs under sub-critical mass conditions.